1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine rotor blade with a serpentine flow cooling circuit having additional turn channel cooling features.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
FIG. 1 shows a prior art turbine rotor blade with a three pass aft flowing serpentine flow cooling circuit 11 with trip strips 12 along the walls of each serpentine circuit channel. A tip turn channel 15 is formed at the blade tip that connects the first leg to the second leg of the serpentine flow circuit, and a root turn channel 19 is formed in the blade root that connects the second leg to the third leg.
FIG. 2 shows a detailed close-up view of the tip turn channel 15 in FIG. 1. The tip turn 15 has a wide open tip turn. For a conical blade tip turn design like that is FIGS. 1 and 2, the downstream turn flow area is much greater than the upstream flow area and thus creates a flow separation and recirculation at locations identified as 17 in FIG. 2. A result of this separation and recirculation of the cooling air flow is that an over-temperature is produced at these locations. An over-temperature is like a hot spot on the blade which leads to erosion damage and thus a shortened part life. This is a major issue for industrial gas turbine engine blades, since these engines must be capable of continuous operation for 40,000 hours or more.
Several prior art patents attempt to address this issue of an over-temperature at the blade tip turns. U.S. Pat. No. 5,073,086 issued to Cooper on Dec. 17, 1991 and entitled COOLED AERFOIL BLADE discloses adding extra material downstream of the turn. U.S. Pat. No. 6,439,848 issued to Haehnle et al on Aug. 27, 2002 and entitled DRILLED COOLING AIR OPENINGS IN GAS TURBINE COMPONENTS discloses adding a bleed hole to purge the flow recirculation and incorporate a turning guide vane in the tip turn region. U.S. Pat. No. 6,939,102 issued to Liang on Sep. 6, 2005 and entitled FLOW GUIDE COMPONENT WITH ENHANCED COOLING discloses t at cooling air is pushed outward for cooling the squealer tip floor and corners while a vortex chamber is used in the middle of the tip turn to provide not only cooling of the tip turn but also purge air for the separation area downstream of the tip turn. U.S. Pat. No. 7,217,097 issued to Liang on May 15, 2007 and entitled COOLING SYSTEM WITH INTERNAL FLOW GUIDE WITHIN A TURBINE BLADE OF A TURBINE ENGINE discloses using a guide vane in the separation flow channel to improve both the tip turn and the root turn flows.